A Comparison of Nuclear Thermal to Nuclear Electric Propulsion for Interplanetary Missions

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A Comparison of Nuclear Thermal to Nuclear Electric Propulsion for Interplanetary Missions

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A Comparison of Nuclear Thermal to Nuclear Electric Propulsion for Interplanetary Missions

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A Comparison of Nuclear Thermal to Nuclear Electric Propulsion for Interplanetary Missions

Mike Osenar

Mentor: LtCol Lawrence

- Introduction
- Objective
- Establish parameters
- NTR Design
- NEP Design
- Discussion and Conclusion

- NASA is developing Nuclear Electric Propulsion (NEP) systems for Project Prometheus, a series of interplanetary missions
- What happened to Nuclear Thermal Rocket (NTR) systems? Should NASA only invest in NEP systems?

- Prove the feasibility of different nuclear propulsion systems for interplanetary missions which fit in a single launch vehicle
- Compare NTR and NEP system designs for given missions
Method: take a set of inputs, use a series of calculations and SPAD process along with reasonable design assumptions to design a spacecraft to reach a given ΔV

- Establish ΔV’s and flight times for both NEP and NTR systems to Jupiter and Pluto
- Determine launch vehicle payload restrictions
- Obtain design points – inert mass fractions based on thruster specific impulses

- NEP ΔV’s and flight times based on AIAA 2002-4729 – low thrust gravity assist trajectories
- NTR data derived from NEP data

- Relationship between NEP ΔV/TOF and NTR ΔV/TOF
- Table shows that NTR has same TOF for 50% of the ΔV
- NTR numbers based on AIAA 1992-3778

Ariane 5 Payload Specifications

Design points established from Dumbkopff charts

Size system so that it meets 3 specifications

- Under max payload mass
- Fits in payload fairing
- Reaches required ΔV

Inputs from Dumbkopff: finert, ΔV

Assumptions

Po = 7 MPa

Isp = 1000 s – hydrogen

Tc = 3200 K

T/W = .3 – experimented, balance between high thrust short burn time and low reactor mass (low power)

- Equations for basic parameters

Subsystem Sizing (note: volume constraint height)

Payload

1000 kg to Jupiter, 500 to Pluto

based on densities of actual space mission

sized as 2 m tall cylinder

Tank

biggest part – hydrogen has low density

Turbo Pump Feed System

Nuclear Reactor

Radiation Shield

standard SPAD design – 18 cm Be, 5 cm W, 5 cm LiH2

Nozzle

Columbium, designed to be ideally expanded in space (ε=100)

Miscellaneous

Avionics

Reactor containment vessel

Attitude thrusters

Structural mass

Payload

Propellant Tank

Pump

Shield

Reactor

Nozzle

Achievable ΔV verified with Rocket Equation

Vehicle height determined by stacking parts according to Figure

Final Results of NTR Design

Size system so that it meets 2 specifications

- Under max payload mass
- Reaches required ΔV
No size requirement – analysis showed that NEP systems would violate mass constraints before volume – no low-density hydrogen propellant

Power Source

- Nuclear Reactors (P>6 kWe)
- Critical reactors designed as small as 6 kWe

- Radioisotope Thermoelectric Generators (RTG) (P<6 kWe)
- Solar?

- Solar Power proportional to inverse square of distance from sun
- to receive power equal to 1 m2 solar panel in earth orbit, would need 27 m2 panel at Jupiter and 1562 m2 panel at Pluto
- does not factor in degradation – significant for long lifetimes
- engineering, GNC concerns with huge solar array
- mass too much

- Thrusters based on actual designed thrusters from SPAD
- Baselines used: T6, XIPS-25, RIT-XT
- Design allowed thrusters to be clustered in groups of up to 3 – proven to work, increases force and power appropriately

- Use NTR equations for propellant mass, thrust, mass flow and power
- NEP equations:

Subsystem Design

- Power system
- Propellant tank
- Thruster mass
- Power conditioning mass
- Other mass (structural, feed systems, avionics, etc.)

NEP Design Results

- Overall, ΔV’s were low – real science mission would need higher ΔV to capture orbit of planet, maneuver
- Accurate data on EP trajectories was desired over ΔV’s for realistic missions

NTR Design

- Almost failed Pluto design – tank volume
- High thrust, impulsive burn more reliable – operates for short time
- Much less efficient then NEP
- Other applications? launch vehicle, human Mars exploration

NEP Design

- Low thrust, long trip times
- Lifetime analysis – electric thrusters tested to 3.5 years – less than Jupiter TOF
- Space Nuclear reactors require extensive testing

- Testing – extensive testing needed for either system – facilities, money needed to test for operational lifetime
- Safety – perennial concern with nuclear systems, real hazards to be considered
- Radiological hazard – higher with NEP (low power but long burn time), must be addressed for either system

- NASA probably right to go with NEP for interplanetary missions
- Much stands between now and operational nuclear propulsion system
- Much to be gained from nuclear propulsion technology

Questions?